The present technology relates generally to a cooling system and method for a gas turbine engine and, in particular, to a system and method of cooling a hot portion of a gas turbine engine.
Gas turbine engines (such as turbojet engines, bypass turbofan engines, turboprop engines, turboshaft engines, etc.) may be used to power flight vehicles (such as planes, helicopters, and missiles, etc.) and may also be used to power ships tanks, electric power generators, pipeline pumping apparatus, etc. For purposes of illustration, the present technology will be described with respect to an aircraft bypass turbofan gas turbine engine. However, it is understood that the present technology is equally applicable to other types and/or uses of gas turbine engines.
Referring to FIG. 9, a gas turbine engine assembly 10 includes a core engine 15 having, in serial flow relationship, a high pressure compressor 18 (also called a core compressor) to compress the airflow entering the core engine 15, a combustor 20 (also called a combustion chamber) in which a mixture of fuel and the compressed air is burned to generate a propulsive gas flow, and a high pressure turbine 22 which is rotated by the propulsive gas flow and which is connected by a larger diameter shaft to drive the high pressure compressor 18. A typical aircraft bypass gas turbine engine adds a low pressure turbine 24 (located aft of the high pressure turbine) which is connected by a smaller diameter coaxial shaft to drive a front fan 14 (located forward of the high pressure compressor) which is surrounded by a nacelle 32 and which may also drive a low pressure compressor 16 (located between the front fan 14 and the high pressure compressor 18). The low pressure compressor 16 sometimes is called a booster compressor or simply a booster. It is understood that the term “compressor” includes, without limitation, high pressure compressors and low pressure compressors. A flow splitter 17, located between the fan 14 and the first (usually the low pressure) compressor, separates the air which exits the fan 14 into a core engine airflow and a surrounding bypass airflow. The bypass airflow from the fan exits the fan bypass duct 40 to provide most of the engine thrust for the aircraft. Some of the engine thrust comes from the core engine airflow after it flows through the low and high pressure compressors 16, 18 to the combustor 20 and is expanded through the high and low pressure turbines 22, 24 and accelerated out of the exhaust nozzle.
Aircraft bypass turbofan gas turbine engines are designed to operate at high temperatures to maximize engine thrust. Cooling of engine hot section components (such as the combustor, the high pressure turbine, the low pressure turbine, and the like) is necessary because of the thermal “redline” limitations of the materials used in the construction of such components. Typically, such cooling of a portion of the engine is accomplished by ducting (also called “bleeding”) cooler air from the high and/or low pressure compressors to those engine components which require such cooling. Unfortunately, the relatively low pressure and hot temperature of the compressor air limits its ability to be used to cool such engine components.
In service, gas turbine aircraft engines are subject to a wide range of operating conditions such as high and low altitudes, high and low temperatures, and high and low speed airflows over, around, and through the engine. Even during a single flight, the aircraft, its engine(s), and engine control components may experience low speed, low altitude, and high temperature conditions during taxi, takeoff, and landing operations, as well as high speed, high altitude, and low temperature conditions during the cruise portion of the flight.
To manage the operating temperatures of the engine components, ventilation is often provided to direct air which is cooler than the components onto the components to carry heat away and maintain the temperature of the component at a satisfactory operating level. However, cooling needs often vary greatly during the course of a flight or operating session. For example, a much greater degree of cooling may be needed on a hot day during ground operations at engine idle power settings than at high altitude during cruise conditions and high power settings.
The technology described herein relates generally to gas turbine engines, and more particularly, to a system and method for cooling engine control components for such engines.